This invention relates generally to gas turbine engine rotor blades, and, more specifically, to blades and bladed disk assemblies of fan and compressor sections thereof.
Bladed disk assemblies, i.e., discrete blades having dovetails mounted in complementary shaped slots in a rotor disk, are well known in the art. Blisk assemblies having integral blades and disks i.e. bl(ade) plus integral (d)isk equals "blisk", see for example, U.S. Pat. No. 4,363,602-J. R. Martin, entitled "Composite Air Foil and Disc Assembly," are also well known in the art. The use of a blisk assembly over a bladed disk assembly provides many benefits including increased structural strength and improved aerodynamic performance. In particular, a blisk can be designed for obtaining a relatively low radius ratio, defined as the inlet hub radius divided by the blade tip radius, having values less than about 0.5, and relatively high blade root solidity, defined as the root chord length divided by the distance between adjacent blades, having values greater than about 2.3 for obtaining significant improvements in aerodynamic performance.
For example, a development engine of the General Electric Company includes a stage 1 titanium compressor blisk having a radius ratio of about 0.42 and a solidity of about 3.1. Although this blisk provides substantial performance benefits, it is deemed desirable to have replaceable blades for more easily repairing any foreign object damage thereto. Furthermore, the use of conventional blade steel material is also deemed desirable for reducing costs.
However, experience has shown that conventional steel-bladed disk assemblies are limited to radius ratios greater than about 0.5 and solidity less than about 2.2 due to life and strength considerations including low-cycle fatigue (LCF) and high-cycle fatigue (HCF). It should be appreciated that for any given compressor stage, the number and size of the blades needed for performing the required amount of work is generally a fixed requirement. With this given number of blades, it will be appreciated that for obtaining reduced radius ratios to improve aerodynamic performance, the outer perimeter of the disk must be correspondingly reduced, thusly providing less circumferential space for mounting the blades thereto and thereby increasing solidity.
Accordingly, smaller shank and dovetail portions of the blade are required due to the physical limitations of the decreased circumference for low radius ratio applications. However, inasmuch as the size of the airfoil portion of the blade does not basically change, the required smaller conventional dovetail and shank are structurally inadequate for suitably mounting the blade to the disk. For example, such a conventional shank and dovetail would be relatively more flexible and have less load transfer surface areas thus leading to undesirable LCF and HCF problems in the dovetail and disk assembly.
Accordingly, it is an object of the present invention to provide a new and improved bladed disk assembly.
Another object of the present invention is to provide a bladed disk assembly which is interchangeable with a blisk assembly having a relatively low radius ratio and relatively high solidity.
Another object of the present invention is to provide a new and improved rotor blade having an improved shank portion.
Another object of the present invention is to provide an improved rotor blade having shank and dovetail portions which are relatively lighter than conventional ones while maintaining acceptable bending stiffness and load-carrying ability.
Another object of the present invention is to provide an improved rotor blade having reduced weight to minimize stress concentration effects.